1. Field
The present disclosure generally relates to composite laminates, such as those used in aircraft, and deals more particularly with a fiber reinforced resin laminate having cross-plies with optimized fiber orientations.
2. Background
Fiber reinforced resin laminates, such as carbon fiber reinforced plastics (CFRP), are widely used in aerospace and other applications because of their favorable strength-to-weight ratio. These composite laminates may be fabricated by laying up multiple plies of unidirectional reinforcing fibers held in a resin matrix, commonly known as prepreg. The plies in the layup may have differing fiber orientations arranged in an order that results in the required laminate strength and stiffness for a particular application. For example, in aircraft skins, the composite laminate may comprise groups of plies respectively having reinforcing fibers oriented at 0, +45, −45 and 90 degrees relative to a reference axis, with the majority of the plies being +/−45 degree plies. Although the number of plies in the laminate may vary at different locations along the wing, the angular orientation of the plies may be substantially constant over the length of the wing, and therefore not fully optimized to match performance requirements at individual locations on the wing. The use of these differing ply orientations allow the composite laminate structure to better resist bending, shear, torsional and bearing loads for a given application, but the use of constant orientations of the plies may result in a skin that is heavier than desired for a particular application.
Reducing the weight of composite laminate structures used in aircraft applications may improve the operating efficiency of an aircraft. The weight of such structures may depend, at least in part, on the number of plies in the laminate, which in turn may be determined by the strength and stiffness requirements for the particular application.
Accordingly, there is a need for a composite laminate that employs a reduced number of total plies while retaining the required laminate strength, rigidity and resistance to splitting and crack propagation. There is also a need for a composite laminate aircraft skin that optimizes the balance between the skin's bending and torsional strength and stiffness by using cross-plies that have optimized fiber orientations. Further, there is a need for a composite laminate skin exhibiting reduced weight through variation of the ply orientation over the wing length.